EP0558904A1 - Aircraft wing with supercritical profile - Google Patents

Aircraft wing with supercritical profile Download PDF

Info

Publication number
EP0558904A1
EP0558904A1 EP93100983A EP93100983A EP0558904A1 EP 0558904 A1 EP0558904 A1 EP 0558904A1 EP 93100983 A EP93100983 A EP 93100983A EP 93100983 A EP93100983 A EP 93100983A EP 0558904 A1 EP0558904 A1 EP 0558904A1
Authority
EP
European Patent Office
Prior art keywords
wing
gap
compensation chamber
ventilation device
wall strip
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP93100983A
Other languages
German (de)
French (fr)
Other versions
EP0558904B1 (en
Inventor
Rainer Prof. Dr. Bohning
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Airbus Operations GmbH
Original Assignee
Deutsche Aerospace AG
Airbus Operations GmbH
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Deutsche Aerospace AG, Airbus Operations GmbH filed Critical Deutsche Aerospace AG
Publication of EP0558904A1 publication Critical patent/EP0558904A1/en
Application granted granted Critical
Publication of EP0558904B1 publication Critical patent/EP0558904B1/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
    • B64C23/04—Influencing air flow over aircraft surfaces, not otherwise provided for by generating shock waves
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C21/00—Influencing air flow over aircraft surfaces by affecting boundary layer flow
    • B64C21/02—Influencing air flow over aircraft surfaces by affecting boundary layer flow by use of slot, ducts, porous areas or the like
    • B64C21/025—Influencing air flow over aircraft surfaces by affecting boundary layer flow by use of slot, ducts, porous areas or the like for simultaneous blowing and sucking
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C2230/00—Boundary layer controls
    • B64C2230/06—Boundary layer controls by explicitly adjusting fluid flow, e.g. by using valves, variable aperture or slot areas, variable pump action or variable fluid pressure
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C2230/00—Boundary layer controls
    • B64C2230/20—Boundary layer controls by passively inducing fluid flow, e.g. by means of a pressure difference between both ends of a slot or duct
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C2230/00—Boundary layer controls
    • B64C2230/22—Boundary layer controls by using a surface having multiple apertures of relatively small openings other than slots
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/10—Drag reduction

Definitions

  • the invention relates to a wing for aircraft with supercritical profiling, which has on its top a ventilation device running in the direction of the span to influence the boundary layer in the region of the compression joint, which consists of a covered with a perforated wall strip and extending on both sides of the compression joint Compensation chamber exists.
  • This method is understood to mean self-regulating blowing and suctioning of flow medium through a porous wall piece in the joint area with the compensation chamber underneath.
  • An increase in pressure across the impact induces a ventilation flow through the porous surface, with part of the boundary layer being sucked off behind the impact and blown out before the impact.
  • Feedback with the external flow with a suitable position and width of the compensation chamber, has a favorable influence on the impact boundary layer interference and thus an improvement in the profile properties.
  • the passage through the wall as proposed in DE-C-33 18 413, can be through a perforation with vertical or oblique holes. This method aims to reduce the wave resistance with the smallest possible increase in the viscous resistance.
  • the invention is based on a supercritical aerofoil with a passive ventilation device arranged on the top in the span direction and has the task of improving the lift-resistance ratio outside the design range (“off-design area”).
  • This object is achieved according to the invention in that the front end of the compensation chamber has a gap-shaped outlet for blowing out the ventilation medium in the direction of the flow.
  • the ventilation medium is blown out tangentially to the wall in the flow direction before the impact, which can be achieved with simple technical means.
  • the ventilation device consists of a combination of a gap and a downstream perforated wall strip with a compensation chamber underneath that connects the two elements.
  • the gap and wall strips extend in the wingspan.
  • a self-regulating equalizing flow is therefore stimulated between the subsonic field and the supersonic field, with part of the boundary layer medium being sucked off behind the impact and being fed tangentially to the incoming boundary layer at a distance from the impact through a gap.
  • the impact is subsequently shifted to the rear end of the chamber and fixed there.
  • the supply of kinetic energy through the gap has a positive effect on the character of the boundary layer.
  • the thickness of the impulse loss and the thickness of the displacement are significantly reduced and the wall shear stress is increased.
  • the speed profile becomes considerably more complete, especially near the wall, which counteracts the tendency of the shock-induced detachment that would otherwise occur. Since tangential blowing, unlike vertical or oblique blowing, does not induce any additional loss of momentum, a higher mass flow through the compensation chamber can be permitted, which increases the ventilation effect.
  • the flight performance under off-design conditions is not only improved, but the off-design area itself is expanded as the beginning of the wing shaking is shifted to higher lift values.
  • the method proposed here has the advantage of being able to do without energy-consuming measures.
  • the proposed method is not only useful for transonic hydrofoils, but can in principle be used wherever shock waves interfere with boundary layers. It is also possible to close the gap if necessary and to prevent the flow through the perforation with the help of a screen on the inside of the wall strip.
  • Fig. 1 is a plan view of a wing 11 flowed with a supercritical Mach number M ⁇ with perforated wall strips 13 and upstream gap plate 12.
  • Wall strips 13 and split plate 12 are attached in such a way that the compression shock in the non-ventilated case is approximately above the center of the wall strip 13 would stand, ie with modern wings 11 at approx. 50 - 70% of the wing depth.
  • the ventilation arrangement 10 is in an optimal position, this results in a positive ventilation effect in the entire off-design area.
  • the ventilation device 10 does not extend over the entire wing span.
  • the compensation chamber 14 underneath is divided in a suitable manner in the span direction in order to avoid pressure compensation in this direction.
  • the perforation of the wall strip 13 can have a hexagonal hole arrangement with a permeability of 4-20%.
  • Fig. 2 shows a section through the wing 11 in the direction of flow. At the top of this profile, a local supersonic area with a final shock can be seen. In the butt region of the wing 11, the compensation chamber 14 is let into the wall, the height of which is of minor importance since the ventilation induces only a relatively small mass flow through the chamber 14. The height is not shown to scale here. In practice, it can be approx. 2 boundary layer thicknesses. The length of the compensation chamber 14 or the distance between the gap 15 and the downstream chamber end can be 5-20% of the profile depth.
  • the compensation chamber 14 is covered with a wall strip 13 which is partially or completely perforated. In both cases, the rear area downstream of the joint is perforated up to the chamber end 16. Upstream, the perforation can extend up to a few boundary layer thicknesses before the impact or even into the gap 15.
  • the flow medium comes from areas of higher pressure (A) through the perforation into the compensation chamber 14 and flows for the most part from the gap 15 into areas of low pressure (B).
  • A areas of higher pressure
  • B low pressure
  • part of the medium can already flow out between the joint and the gap. In this case there is a mixing of the tangentially and vertically flowing Medium.
  • This version thus represents a combination of the method proposed here and that proposed in DE 33 18 413. Care must be taken when designing the edge at the exit of the gap 15 so that the boundary layer is not burdened by detachment vortices which may be induced here.
  • FIGS. 4 to 8 relate to a numerical simulation of the flow around a specific supercritical airfoil profile. Cases a) without a ventilation device are compared to cases b) with a ventilation device.
  • the lift coefficient C a is 0.6 in each case in the representations according to FIGS. 4 to 7.
  • FIG. 4 shows calculated isotachic fields and FIG. 5 the associated wall pressure profiles.
  • a comparison of cases a) and b) shows that the impact is shifted downstream near the end of the chamber (FIG. 5b), while the angle of attack decreases to 1 o for the same lift coefficient. It can be concluded from this that with a constant angle of attack in case b) a considerable increase in lift compared to the non-ventilated case can be achieved. It can be seen from the diagrams for the displacement thickness and the pulse loss thickness (FIG. 6) that both boundary layer sizes are drastically reduced with ventilation. The lower viscous resistance results from the reduction in the pulse loss thickness. The course of the coefficient of friction c f (FIG. 7) indicates a significantly lower tendency to detach compared to the non-ventilated case.
  • the graphs in Fig. 8 show the polar for buoyancy, total drag, wave drag and viscous drag. It can be seen that in the case of profiles with a ventilation device according to the invention, the viscous resistance is drastically reduced, while the wave resistance increases, but the sum of both resistances, namely the total resistance, decreases significantly. In addition, ventilation with tangential blowing leads to a considerable increase in lift.

Abstract

The invention relates to an aircraft wing (11) having a supercritical profile, which has a ventilation device, running in the wing span direction, on the top. The ventilation device consists of a compensating chamber (14), which is covered by a perforated wall strip (13), extends on both sides of the compression shock and whose front end has an outlet (15), in the form of a gap, for blowing the ventilation medium out in the flow direction. <IMAGE>

Description

Die Erfindung bezieht sich auf einen Tragflügel für Luftfahrzeuge mit überkritischer Profilierung, welcher an seiner Oberseite eine in Richtung der Spannweite verlaufende Ventilationsvorrichtung zur Beeinflussung der Grenzschicht im Bereich des Verdichtungsstoßes aufweist, die aus einer mit einem perforierten Wandstreifen abgedeckten und zu beiden Seiten des Verdichtungsstoßes sich erstreckenden Ausgleichskammer besteht.The invention relates to a wing for aircraft with supercritical profiling, which has on its top a ventilation device running in the direction of the span to influence the boundary layer in the region of the compression joint, which consists of a covered with a perforated wall strip and extending on both sides of the compression joint Compensation chamber exists.

Ein derartiger Tragflügel ist in der DE-C-33 18 413 ausführlich beschrieben. Wie in dieser Druckschrift angegeben, stellt sich bei einer transsonischen Umströmung solcher Tragflügel auf der Oberseite im Bereich des Dickenmaximums ein lokales Überschallgebiet ein und für einen bestimmten Auslegungszustand kann der Übergang vom Überschall zum Unterschall nahezu stoßfrei durch geeignete Profilierung realisiert werden. Bei geringfügigen Änderungen der Anströmbedingungen (sog. "off-design Bereich") wird das Überschallgebiet aber in der Regel durch einen Verdichtungsstoß stromabwärts abgeschlossen, welcher eine beträchtliche Widerstandsquelle darstellt. Aufgrund der Wechselwirkung des Stoßes mit der wandnahen Reibungsgrenzschicht kann es je nach Stoßstärke zur Strömungsablösung und als Folge davon zum höchst unerwünschten Tragflügelschütteln kommen.Such a wing is described in detail in DE-C-33 18 413. As indicated in this publication, a transonic flow around such wings on the upper side in the area of the thickness maximum results in a local supersonic area and for a certain design state the Transition from supersonic to subsonic can be realized almost without bumps by suitable profiling In the event of slight changes in the inflow conditions (so-called "off-design area"), however, the supersonic area is usually closed off by a compression shock downstream, which represents a considerable source of resistance. Due to the interaction of the impact with the friction boundary layer close to the wall, depending on the impact strength, flow separation and, as a result, highly undesirable wing shaking can occur.

Durch Beeinflussung des Stoß-Grenzschicht-Interferenzgebietes ist es möglich diesen negativen Auswirkungen auf einfache Weise passiv entgegenwirken. Man versteht unter dieser Methode ein sich selbstregulierendes Ausblasen und Absaugen von Strömungsmedium durch ein poröses Wandstück im Stoßbereich mit darunterliegender Ausgleichskammer. Ein Druckanstieg über den Stoß hinweg induziert hierbei eine Ventilationsströmung durch die poröse Oberfläche, wobei ein Teil der Grenzschicht hinter dem Stoß abgesaugt und vor dem Stoß wieder ausgeblasen wird. Durch Rückkopplung mit der Außenströmung kommt es bei geeigneter Lage und Breite der Ausgleichskammer zu einer günstigen Beeinflussung der Stoß-Grenzschicht-Interferenz und damit zu einer Verbesserung der Profileigenschaften. Der Durchlaß durch die Wand kann, wie in der DE-C-33 18 413 vorgeschlagen, durch eine Perforation mit senkrechten oder schrägen Bohrungen erfolgen. Diese Methode zielt auf eine Reduktion des Wellenwiderstandes bei möglichst geringem Anstieg des viskosen Widerstandes ab.By influencing the impact boundary layer interference area, it is possible to passively counteract these negative effects in a simple manner. This method is understood to mean self-regulating blowing and suctioning of flow medium through a porous wall piece in the joint area with the compensation chamber underneath. An increase in pressure across the impact induces a ventilation flow through the porous surface, with part of the boundary layer being sucked off behind the impact and blown out before the impact. Feedback with the external flow, with a suitable position and width of the compensation chamber, has a favorable influence on the impact boundary layer interference and thus an improvement in the profile properties. The passage through the wall, as proposed in DE-C-33 18 413, can be through a perforation with vertical or oblique holes. This method aims to reduce the wave resistance with the smallest possible increase in the viscous resistance.

Die Erfindung geht von einem überkritischen Tragflügel mit einer auf der Oberseite angeordneten in Spannweitenrichtung verlaufenden passiven Ventilationsvorrichtung aus und hat die Aufgabe das Auftriebs-Widerstandsverhältnis außerhalb des Auslegebereiches ("Off-design Bereich") zu verbessern. Diese Aufgabe ist gemäß der Erfindung dadurch gelöst, daß das vordere Ende der Ausgleichskammer einen spaltförmigen Austritt zum Ausblasen des Ventilationsmediums in Richtung der Anströmung aufweist.The invention is based on a supercritical aerofoil with a passive ventilation device arranged on the top in the span direction and has the task of improving the lift-resistance ratio outside the design range (“off-design area”). This object is achieved according to the invention in that the front end of the compensation chamber has a gap-shaped outlet for blowing out the ventilation medium in the direction of the flow.

Bei der erfindungsgemäßen Ventilationsvorrichtung wird das Ventilationsmedium vor dem Stoß tangential zur Wand in Strömungsrichtung ausgeblasen, was sich mit einfachen technischen Mitteln verwirklichen läßt. Die Ventialtionsvorrichtung besteht aus einer Kombination von Spalt und stromabwärts angebrachtem perforierten Wandstreifen mit darunterliegender Ausgleichskammer, die beide Elemente verbindet. Spalt und Wandstreifen erstrecken sich in Flügelspannweite. Zwischen dem Unterschallfeld und dem überschallfeld wird daher eine sich selbstregulierende Ausgleichsströmung angeregt, wobei ein Teil des Grenzschichtmediums hinter dem Stoß abgesaugt und in einiger Entfernung vor dem Stoß durch einen Spalt der ankommenden Grenzschicht tangential zugeleitet wird. Dadurch wird in der Folge der Stoß zum hinteren Kammerende verschoben und dort fixiert. Die Zufuhr von kinetischer Energie durch den Spalt wirkt sich hierbei positiv auf den Charakter der Grenzschicht aus. Dabei werden die Impulsverlustdicke und die Verdrängungsdicke deutlich reduziert und die Wandschubspannung erhöht. Das Geschwindigkeitsprofil wird, besonders in Wandnähe, beträchtlich völliger, wodurch der Tendenz der sonst auftretenden stoßinduzierten Ablösung entsprechend entgegengewirkt wird. Da das tangentiale Ausblasen im Gegensatz zum vertikalen oder schrägen Ausblasen keinen zusätzlichen Impulsverlust induziert, kann ein höherer Massenstrom durch die Ausgleichskammer zugelassen werden, was den Ventilationseffekt verstärkt.In the ventilation device according to the invention, the ventilation medium is blown out tangentially to the wall in the flow direction before the impact, which can be achieved with simple technical means. The ventilation device consists of a combination of a gap and a downstream perforated wall strip with a compensation chamber underneath that connects the two elements. The gap and wall strips extend in the wingspan. A self-regulating equalizing flow is therefore stimulated between the subsonic field and the supersonic field, with part of the boundary layer medium being sucked off behind the impact and being fed tangentially to the incoming boundary layer at a distance from the impact through a gap. As a result, the impact is subsequently shifted to the rear end of the chamber and fixed there. The supply of kinetic energy through the gap has a positive effect on the character of the boundary layer. The thickness of the impulse loss and the thickness of the displacement are significantly reduced and the wall shear stress is increased. The speed profile becomes considerably more complete, especially near the wall, which counteracts the tendency of the shock-induced detachment that would otherwise occur. Since tangential blowing, unlike vertical or oblique blowing, does not induce any additional loss of momentum, a higher mass flow through the compensation chamber can be permitted, which increases the ventilation effect.

Für den überkritischen Flügel ergibt sich bei geeigneter Ausgestaltung der Ventilationsvorrichtung insgesamt eine Widerstandsverringerung im "Off-design" Bereich. Zusätzlich kommt es zu einer beträchtlichen Erhöhung des Auftriebs. Die Reduzierung des Gesamtwiderstandes basiert hier auf der durch das tangentiale Ausblasen hervorgerufenen starken Absenkung des viskosen Widerstandes. Der Auftriebsgewinn resultiert aus der Verschiebung des Stoßes stromabwärts.With a suitable design of the ventilation device, there is an overall reduction in resistance in the "off-design" area for the supercritical wing. In addition, there is a significant increase in buoyancy. The reduction in the total resistance is based here on the strong reduction in the viscous resistance caused by the tangential blowing. The gain in lift results from the displacement of the shock downstream.

Durch die Verringerung des Gesamtwiderstandes und gleichzeitige Erhöhung des Auftriebs werden die Flugleistungen unter Off-design Bedingungen nicht nur verbessert, sondern es wird der Off-design Bereich selbst erweitert, da der Beginn des Tragflügelschüttelns zu höheren Auftriebswerten verschoben wird. Gegenüber aktiven Methoden zur Stoßbeeinflussung bietet die hier vorgeschlagene Methode den Vorteil, ohne energieverbrauchende Maßnahmen auszukommen.By reducing the overall resistance and simultaneously increasing the lift, the flight performance under off-design conditions is not only improved, but the off-design area itself is expanded as the beginning of the wing shaking is shifted to higher lift values. Compared to active methods for influencing impacts, the method proposed here has the advantage of being able to do without energy-consuming measures.

Die vorgeschlagene Methode ist nicht nur für transsonische Tragflügel von Nutzen, sondern sie läßt sich prinzipiell überall dort einsetzen, wo Verdichtungsstöße mit Grenzschichten interferieren. Dabei ist es auch möglich, bei Bedarf den Spalt zu verschließen und die Strömung durch die Perforation mit Hilfe einer Blende auf der Innenseite des Wandstreifens zu unterbinden.The proposed method is not only useful for transonic hydrofoils, but can in principle be used wherever shock waves interfere with boundary layers. It is also possible to close the gap if necessary and to prevent the flow through the perforation with the help of a screen on the inside of the wall strip.

Die Erfindung wird anhand der Zeichnung näher erläutert. Es zeigen:

Fig. 1
eine Draufsicht auf einen Ausschnitt eines Luftfahrzeuges mit einem transsonischen Tragflügel,
Fig. 2
einen Querschnitt durch den transsonischen Flügel gemäß Fig. 1,
Fig. 3
ein Stoß-Grenzschicht-Interferenzgebiet mit Ventilationsvorrichtung,
Fig. 4a
ein Isotachenfeld für einen Tragflügel ohne Ventilation,
Fig. 4b
ein Isotachenfeld für einen Tragflügel mit Ventialtion, tangentialem Ausblasen und wandnormalem Absaugen,
Fig. 5
berechnete Druckverläufe am Grenzschichtrand gemäß Fig. 4a, b,
Fig. 6
berechnete Verdrängungsdicken und Impulsverlustdicken gemäß Fig. 4a, b,
Fig. 7
berechnete Reibungsbeiwerte gemäß Fig. 4a, b und
Fig. 8
Diagramme für Auftrieb, Gesamtwiderstand, Wellenwiderstand und viskosen Widerstand.
The invention is explained in more detail with reference to the drawing. Show it:
Fig. 1
2 shows a plan view of a section of an aircraft with a transonic wing,
Fig. 2
2 shows a cross section through the transonic wing according to FIG. 1,
Fig. 3
an impact boundary layer interference area with ventilation device,
Fig. 4a
an isotachian field for a wing without ventilation,
Fig. 4b
an isotachian field for a wing with ventialtion, tangential blowing and suction normal to the wall,
Fig. 5
calculated pressure profiles at the boundary layer edge according to FIGS. 4a, b,
Fig. 6
calculated displacement thicknesses and pulse loss thicknesses according to FIGS. 4a, b,
Fig. 7
calculated coefficients of friction according to FIGS. 4a, b and
Fig. 8
Charts for buoyancy, total drag, wave drag and viscous drag.

In Fig. 1 ist eine Draufsicht auf einen mit einer überkritischen Machzahl M∞ angeströmten Tragflügel 11 mit perforiertem Wandstreifen 13 und vorgelagertem Spaltblech 12 abgebildet. Wandstreifen 13 und Spaltblech 12 sind so angebracht, daß der Verdichtungsstoß im nichtventilierten Fall etwa über der Mitte des Wandstreifens 13 stehen würde, d.h. bei modernen Tragflügeln 11 bei ca. 50 - 70 % der Flügeltiefe. Bei optimaler Lage der Ventilationsanordnung 10 ergibt sich so ein positiver Ventilationseffekt im gesamten Offdesign Bereich. Wie die Zeichnung zeigt, erstreckt sich die Ventilationsvorrichtung 10 nicht über die gesamte Flügelspannweite. Die darunter befindliche Ausgleichskammer 14 ist in Spannweitenrichtung in geeigneter Weise unterteilt, um einen Druckausgleich in dieser Richtung zu vermeiden. Die Perforation des Wandstreifens 13 kann eine hexagonale Lochanordnung aufweisen mit einer Durchlässigkeit von 4 - 20 %.In Fig. 1 is a plan view of a wing 11 flowed with a supercritical Mach number M∞ with perforated wall strips 13 and upstream gap plate 12. Wall strips 13 and split plate 12 are attached in such a way that the compression shock in the non-ventilated case is approximately above the center of the wall strip 13 would stand, ie with modern wings 11 at approx. 50 - 70% of the wing depth. If the ventilation arrangement 10 is in an optimal position, this results in a positive ventilation effect in the entire off-design area. As the drawing shows, the ventilation device 10 does not extend over the entire wing span. The compensation chamber 14 underneath is divided in a suitable manner in the span direction in order to avoid pressure compensation in this direction. The perforation of the wall strip 13 can have a hexagonal hole arrangement with a permeability of 4-20%.

Fig. 2 zeigt einen Schnitt durch den Flügel 11 in Strömungsrichtung. An der Oberseite dieses Profils ist ein lokales überschallgebiet mit abschließendem Verdichtungsstoß zu erkennen. Im Stoßbereich des Tragflügels 11 ist in die Wand die Ausgleichskammer 14 eingelassen, deren Höhe von untergeordneter Bedeutung da die Ventilation nur einen relativ geringen Massenstrom durch die Kammer 14 induziert. Die Höhe ist hier nicht maßstäblich dargestellt. Sie kann in der Praxis ca. 2 Grenzschichtdicken betragen. Die Länge der Ausgleichskammer 14 bzw. der Abstand zwischen Spalt 15 und stromabwärts liegendem Kammerende kann 5 - 20 % der Profiltiefe betragen.Fig. 2 shows a section through the wing 11 in the direction of flow. At the top of this profile, a local supersonic area with a final shock can be seen. In the butt region of the wing 11, the compensation chamber 14 is let into the wall, the height of which is of minor importance since the ventilation induces only a relatively small mass flow through the chamber 14. The height is not shown to scale here. In practice, it can be approx. 2 boundary layer thicknesses. The length of the compensation chamber 14 or the distance between the gap 15 and the downstream chamber end can be 5-20% of the profile depth.

Fig. 3 verdeutlicht schematisch die wesentlichen Details der Ventilationsvorrichtung 10. Die Ausgleichskammer 14 ist mit einem Wandstreifen 13 abgedeckt, der teilweise oder auf der ganzen Länge perforiert ist. In beiden Fällen ist jeweils der hintere, stromabwärts vom Stoß liegende, Bereich bis zum Kammerende 16 perforiert. Stromaufwärts kann sich die Perforation bis zu einigen Grenzschichtdicken vor den Stoß oder sogar bis in den Spalt 15 erstrecken. Das Strömungsmedium gelangt aus Gebieten höheren Druckes (A) durch die Perforation in die Ausgleichskammer 14 und strömt zum größten Teil aus dem Spalt 15 in Gebiete niederen Druckes (B). Ein Teil des Mediums kann je nach Durchlässigkeit der Wand 13 auch schon zwischen Stoß und Spalt ausströmen. In diesem Fall kommt es zu einer Vermischung des tangential und vertikal ausströmenden Mediums. Diese Version stellt somit eine Kombination der hier vorgeschlagenen und der in DE 33 18 413 vorgeschlagenen Methode dar. Bei der Gestaltung der Kante am Ausgang des Spaltes 15 ist Sorgfalt geboten, damit die Grenzschicht nicht durch hier eventuell induzierte Ablösewirbel belastet wird.3 schematically illustrates the essential details of the ventilation device 10. The compensation chamber 14 is covered with a wall strip 13 which is partially or completely perforated. In both cases, the rear area downstream of the joint is perforated up to the chamber end 16. Upstream, the perforation can extend up to a few boundary layer thicknesses before the impact or even into the gap 15. The flow medium comes from areas of higher pressure (A) through the perforation into the compensation chamber 14 and flows for the most part from the gap 15 into areas of low pressure (B). Depending on the permeability of the wall 13, part of the medium can already flow out between the joint and the gap. In this case there is a mixing of the tangentially and vertically flowing Medium. This version thus represents a combination of the method proposed here and that proposed in DE 33 18 413. Care must be taken when designing the edge at the exit of the gap 15 so that the boundary layer is not burdened by detachment vortices which may be induced here.

Die Darstellungen nach Fig. 4 bis Fig. 8 beziehen sich auf eine numerische Simulation der Umströmung eines bestimmten überkritischen Tragflügelprofils. Dabei werden jeweils die Fälle a) ohne Ventilationsvorrichtung den Fällen b) mit Ventilationsvorrichtung gegenübergestellt.The representations according to FIGS. 4 to 8 relate to a numerical simulation of the flow around a specific supercritical airfoil profile. Cases a) without a ventilation device are compared to cases b) with a ventilation device.

Bei dem verwendeten Profil handelt es sich um das Profil LVA-1A bei einer Anströmmachzahl von M∞ = 0,75 und einer Reynoldszahl von Re = 6 · 10⁶. Der Auftriebsbeiwert Ca beträgt bei den Darstellungen nach Fig. 4 bis Fig. 7 jeweils 0,6.The profile used is the LVA-1A profile with an inflow number of M∞ = 0.75 and a Reynolds number of Re = 6 · 10⁶. The lift coefficient C a is 0.6 in each case in the representations according to FIGS. 4 to 7.

Fig. 4 zeigt berechnete Isotachenfelder und Fig. 5 die zugehörigen Wanddruckverläufe. Ein Vergleich der Fälle a) und b) zeigt, daß der Stoß stromabwärts in die Nähe des Kammerendes verschoben wird (Fig. 5b), während der Anstellwinkel bei gleichem Auftriebsbeiwert auf 1o abnimmt. Daraus ist zu schließen, daß bei konstant gehaltenem Anstellwinkel im Fall b) ein beträchtlicher Auftriebsgewinn gegenüber dem nichtventilierten Fall erzielt werden kann. Den Diagrammen für die Verdrängungsdicke und die Impulsverlustdicke (Fig. 6) kann entnommen werden, daß beide Grenzschichtgrößen bei Ventilation drastisch reduziert werden. Aus der Verringerung der Impulsverlustdicke resultiert der geringere viskose Widerstand. Der Verlauf des Reibungsbeiwertes cf (Fig. 7) deutet auf eine wesentlich geringere Tendenz zur Ablösung gegenüber dem nichtventilierten Fall hin.FIG. 4 shows calculated isotachic fields and FIG. 5 the associated wall pressure profiles. A comparison of cases a) and b) shows that the impact is shifted downstream near the end of the chamber (FIG. 5b), while the angle of attack decreases to 1 o for the same lift coefficient. It can be concluded from this that with a constant angle of attack in case b) a considerable increase in lift compared to the non-ventilated case can be achieved. It can be seen from the diagrams for the displacement thickness and the pulse loss thickness (FIG. 6) that both boundary layer sizes are drastically reduced with ventilation. The lower viscous resistance results from the reduction in the pulse loss thickness. The course of the coefficient of friction c f (FIG. 7) indicates a significantly lower tendency to detach compared to the non-ventilated case.

Die Diagramme in Fig. 8 zeigen die Polaren für den Auftrieb, Gesamtwiderstand, Wellenwiderstand und viskosen Widerstand. Es wird ersichtlich, daß bei Profilen mit erfindungsgemäßer Ventilationsvorrichtung der viskose Widerstand drastisch verringert wird, während der Wellenwiderstand zwar zunimmt, aber die Summe aus beiden Widertänden, und zwar der Gesamtwiderstand, deutlich abnimmt. Darüber hinaus führt die Ventilation mit tangentialem Ausblasen zu einem beträchtlichen Auftriebsgewinn.The graphs in Fig. 8 show the polar for buoyancy, total drag, wave drag and viscous drag. It can be seen that in the case of profiles with a ventilation device according to the invention, the viscous resistance is drastically reduced, while the wave resistance increases, but the sum of both resistances, namely the total resistance, decreases significantly. In addition, ventilation with tangential blowing leads to a considerable increase in lift.

Claims (6)

Tragflügel für Luftfahrzeuge mit überkritischer Profilierung, welcher an seiner Oberseite eine in Richtung der Spannweite verlaufende Ventilationsvorrichtung zur Beeinflussung der Grenzschicht im Bereich des Verdichtungsstoßes aufweist, die aus einer mit einem perforierten Wandstreifen abgedeckten und zu beiden Seiten des Verdichtungsstoßes sich erstreckenden Ausgleichskammer besteht, dadurch gekennzeichnet, daß das vordere Ende der Ausgleichskammer (14) einen spaltförmigen Austritt (15) zum Ausblasen des Ventilationsmediums in Richtung der Hauptströmung aufweist.Wing for aircraft with supercritical profiling, which has on its top a ventilation device running in the direction of the wingspan to influence the boundary layer in the area of the compression joint, which consists of a compensation chamber covered with a perforated wall strip and extending on both sides of the compression joint, characterized in that that the front end of the compensation chamber (14) has a gap-shaped outlet (15) for blowing out the ventilation medium in the direction of the main flow. Tragflügel nach Anspruch 1, dadurch gekennzeichnet, daß der spaltförmige Austritt (15) zum stromabwärts liegendem Ende (16) des perforierten Wandstreifens (13) einen Abstand aufweist, der auf einen Wert von 5 % bis 20 % der örtlichen Flügeltiefe bemessen ist, und daß die Mitte dieses Abstandes zwischen 50 % und 70 % der örtlichen Flügeltiefe (Stoßlage) auf der Flügeloberseite angeordnet ist.Wing according to Claim 1, characterized in that the gap-shaped outlet (15) is at a distance from the downstream end (16) of the perforated wall strip (13) which is dimensioned at a value of 5% to 20% of the local wing depth, and in that the middle of this distance is located between 50% and 70% of the local wing depth (abutting position) on the top of the wing. Tragflügel nach einem der Ansprüche 1 oder 2, dadurch gekennzeichnet, daß die aus spaltförmigem Austritt (15) perforiertem Wandstreifen (13) und Ausgleichskammer (14) bestehende Ventilationsvorrichtung (10) sich nur über einen Teil der Flügelspannweite erstreckt.Wing according to one of claims 1 or 2, characterized in that the wall strip (13) and compensation chamber (14) perforated from the gap-shaped outlet (15) and the compensation chamber (14) extends only over part of the wing span. Tragflügel nach einem der Ansprüche 1 bis 3, dadurch gekennzeichnet, daß die Ausgleichskammer (14) mit spaltförmigem Austritt (15) und Ventilationseinrichtung (10) in Richtung der Flügelspannweite durch Zwischenwände unterteilt ist.A wing according to one of claims 1 to 3, characterized in that the compensation chamber (14) with a gap-shaped outlet (15) and ventilation device (10) is divided in the direction of the wing span by partitions. Tragflügel nach einem der Ansprüche 1 bis 4, dadurch gekennzeichnet, daß der spaltförmige Austritt (15) in seiner Höhe und die Perforation des Wandstreifens (13) durch auf der Innenseite vorgesehene Blenden veränderbar sind.Wing according to one of claims 1 to 4, characterized in that the height of the gap-shaped outlet (15) and the perforation of the wall strip (13) can be changed by means of panels provided on the inside. Tragflügel nach einem der Ansprüche 1 bis 5, dadurch gekennzeichnet, daß die Ventilationsvorrichtung (10) mit Austrittspalt (15) und perforiertem Wandstreifen (13) an Orten einsetzbar ist, wo Verdichtungsstöße mit Grenzschichten interferieren.Wing according to one of claims 1 to 5, characterized in that the ventilation device (10) with an outlet gap (15) and perforated wall strip (13) can be used at locations where compression surges interfere with boundary layers.
EP93100983A 1992-03-06 1993-01-22 Aircraft wing with supercritical profile Expired - Lifetime EP0558904B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE4207103 1992-03-06
DE4207103A DE4207103C1 (en) 1992-03-06 1992-03-06

Publications (2)

Publication Number Publication Date
EP0558904A1 true EP0558904A1 (en) 1993-09-08
EP0558904B1 EP0558904B1 (en) 1995-11-22

Family

ID=6453387

Family Applications (1)

Application Number Title Priority Date Filing Date
EP93100983A Expired - Lifetime EP0558904B1 (en) 1992-03-06 1993-01-22 Aircraft wing with supercritical profile

Country Status (3)

Country Link
US (1) US5335885A (en)
EP (1) EP0558904B1 (en)
DE (1) DE4207103C1 (en)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0679572A2 (en) * 1994-04-28 1995-11-02 British Aerospace Public Limited Company Laminar flow skin
US5613649A (en) * 1994-07-21 1997-03-25 United Technologies Corporation Airfoil noise control
US6796533B2 (en) * 2001-03-26 2004-09-28 Auburn University Method and apparatus for boundary layer reattachment using piezoelectric synthetic jet actuators
DE10332665B3 (en) * 2003-07-18 2005-01-20 Deutsches Zentrum für Luft- und Raumfahrt e.V. Method for reducing of wave drag entails use of groups each with number of small spoilers arranged on surface of aerofoil and extended in flow direction and spaced apart at right angles to flow direction
WO2010099967A1 (en) * 2009-03-04 2010-09-10 Airbus Operations Gmbh Wing of an aircraft and assembly of a wing comprising a device for influencing a flow
WO2010084025A3 (en) * 2009-01-26 2011-12-08 Airbus Operations Gmbh High-lift flap, arrangement of a high-lift flap together with a device for influencing the flow on the same and aircraft comprising said arrangement
DE102019130596A1 (en) * 2019-11-13 2021-05-20 Deutsches Zentrum für Luft- und Raumfahrt e.V. Method for suppressing aeroelastic instabilities on transonic operating aircraft and aircraft with devices for carrying out the method by means of direct intervention in the aerodynamic boundary layer

Families Citing this family (28)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE4422152C2 (en) * 1994-06-27 2000-02-03 Daimler Chrysler Aerospace Method and arrangement for optimizing the aerodynamic effect of a wing
DE4442319C1 (en) * 1994-11-29 1996-05-09 Deutsche Forsch Luft Raumfahrt Aircraft wing designed for speeds in high subsonic range
GB2296306B (en) * 1994-12-05 1999-05-26 Short Brothers Plc Aeroengine low drag nacelle structure
GB9424495D0 (en) * 1994-12-05 1995-01-25 Short Brothers Plc Aerodynamic low drag structure
US5875998A (en) * 1996-02-05 1999-03-02 Daimler-Benz Aerospace Airbus Gmbh Method and apparatus for optimizing the aerodynamic effect of an airfoil
US6821090B1 (en) 1997-06-05 2004-11-23 Mcdonnell Douglas Helicopter Company Gust alleviation/flutter suppression device
US6543719B1 (en) 1997-06-05 2003-04-08 Mcdonnell Douglas Helicopter Co. Oscillating air jets for implementing blade variable twist, enhancing engine and blade efficiency, and reducing drag, vibration, download and ir signature
US20030201367A1 (en) * 2002-04-24 2003-10-30 Saiz Manuel Munoz Process for the recovery of the energy from the air in pressurised areas of aircraft
US7475853B2 (en) * 2002-06-21 2009-01-13 Darko Segota Method and system for regulating external fluid flow over an object's surface, and particularly a wing and diffuser
US6866233B2 (en) * 2003-01-03 2005-03-15 Orbital Research Inc. Reconfigurable porous technology for fluid flow control and method of controlling flow
DE10305973B3 (en) * 2003-02-13 2004-05-06 Deutsches Zentrum für Luft- und Raumfahrt e.V. Loss reduction device for compression shock includes at least one shock-inducing body spaced out from surface in supersonic flow region
US8485476B2 (en) * 2004-08-20 2013-07-16 University Of Miami Discrete co-flow jet (DCFJ) airfoil
WO2008054518A2 (en) * 2006-04-28 2008-05-08 University Of Miami Emissionless silent and ultra-efficient airplane using cfj airfoil
US8262031B2 (en) * 2004-08-20 2012-09-11 University Of Miami Co-flow jet aircraft
US6994297B1 (en) * 2004-10-27 2006-02-07 The Boeing Company Method and apparatus for controlling a vehicle
US9908617B2 (en) * 2005-07-25 2018-03-06 The Boeing Company Active flow control for transonic flight
US20070029403A1 (en) * 2005-07-25 2007-02-08 The Boeing Company Dual point active flow control system for controlling air vehicle attitude during transonic flight
US9541106B1 (en) * 2007-01-03 2017-01-10 Orbitel Research Inc. Plasma optimized aerostructures for efficient flow control
US8974177B2 (en) 2010-09-28 2015-03-10 United Technologies Corporation Nacelle with porous surfaces
EP3064699B1 (en) 2015-03-03 2018-04-25 Kochs GmbH Window or door, and method for the production of same
DE202015101331U1 (en) 2015-03-03 2015-04-15 Kochs Gmbh Window or door
RU2615251C1 (en) * 2015-11-12 2017-04-04 Федеральное государственное унитарное предприятие "Центральный аэрогидродинамический институт имени профессора Н.Е. Жуковского" (ФГУП "ЦАГИ") Method to weaken shock-induced separation under shockwave boundary layer interaction
US10106246B2 (en) 2016-06-10 2018-10-23 Coflow Jet, LLC Fluid systems that include a co-flow jet
US10315754B2 (en) 2016-06-10 2019-06-11 Coflow Jet, LLC Fluid systems that include a co-flow jet
US10683076B2 (en) 2017-10-31 2020-06-16 Coflow Jet, LLC Fluid systems that include a co-flow jet
US11293293B2 (en) 2018-01-22 2022-04-05 Coflow Jet, LLC Turbomachines that include a casing treatment
US11111025B2 (en) 2018-06-22 2021-09-07 Coflow Jet, LLC Fluid systems that prevent the formation of ice
WO2021016321A1 (en) 2019-07-23 2021-01-28 Gecheng Zha Fluid systems and methods that address flow separation

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR929942A (en) * 1946-06-17 1948-01-12 Sncase Method and devices for controlling transonic flows
FR2489244A1 (en) * 1980-08-28 1982-03-05 Colignon Bernard Pressure equalising vent for airfoil - has vents to reduce boundary layer at rear of air foil
US4522360A (en) * 1983-04-27 1985-06-11 Rensselaer Polytechnic Institute Passive drag control of airfoils at transonic speeds
GB2187261A (en) * 1986-01-06 1987-09-03 Secr Defence Controlling boundary layer

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1913644A (en) * 1931-11-18 1933-06-13 Edward A Stalker Means of energizing the boundary layer on aircraft parts
US3915412A (en) * 1972-05-09 1975-10-28 Robert C Tibbs Airfoil construction
US4137008A (en) * 1977-09-21 1979-01-30 The United States Of America As Represented By The Secretary Of The Navy Adjustable blowing slot for circulation control airfoil
DE2949133C2 (en) * 1979-12-06 1984-11-15 Messerschmitt-Bölkow-Blohm GmbH, 8012 Ottobrunn Airfoil with supercritical profiling
US4457480A (en) * 1981-02-25 1984-07-03 The United States Of America As Represented By The Secretary Of The Navy Mono-element combined supercritical high lift airfoil
DE3318413A1 (en) * 1983-05-20 1984-11-22 Messerschmitt-Bölkow-Blohm GmbH, 8012 Ottobrunn WING FOR AIRCRAFT
DE3811616C1 (en) * 1988-04-07 1989-07-27 Messerschmitt-Boelkow-Blohm Gmbh, 8012 Ottobrunn, De
US5167387A (en) * 1991-07-25 1992-12-01 Vigyan, Inc. Porous airfoil and process

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR929942A (en) * 1946-06-17 1948-01-12 Sncase Method and devices for controlling transonic flows
FR2489244A1 (en) * 1980-08-28 1982-03-05 Colignon Bernard Pressure equalising vent for airfoil - has vents to reduce boundary layer at rear of air foil
US4522360A (en) * 1983-04-27 1985-06-11 Rensselaer Polytechnic Institute Passive drag control of airfoils at transonic speeds
GB2187261A (en) * 1986-01-06 1987-09-03 Secr Defence Controlling boundary layer

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0679572A2 (en) * 1994-04-28 1995-11-02 British Aerospace Public Limited Company Laminar flow skin
EP0679572A3 (en) * 1994-04-28 1996-04-17 British Aerospace Laminar flow skin.
US5848768A (en) * 1994-04-28 1998-12-15 British Aerospace Plc Laminar flow skin
US5613649A (en) * 1994-07-21 1997-03-25 United Technologies Corporation Airfoil noise control
US6796533B2 (en) * 2001-03-26 2004-09-28 Auburn University Method and apparatus for boundary layer reattachment using piezoelectric synthetic jet actuators
DE10332665B3 (en) * 2003-07-18 2005-01-20 Deutsches Zentrum für Luft- und Raumfahrt e.V. Method for reducing of wave drag entails use of groups each with number of small spoilers arranged on surface of aerofoil and extended in flow direction and spaced apart at right angles to flow direction
WO2010084025A3 (en) * 2009-01-26 2011-12-08 Airbus Operations Gmbh High-lift flap, arrangement of a high-lift flap together with a device for influencing the flow on the same and aircraft comprising said arrangement
US10173768B2 (en) 2009-01-26 2019-01-08 Airbus Operations Gmbh High-lift flap, arrangement of a high-lift flap together with a device for influencing the flow on the same and aircraft comprising said arrangement
WO2010099967A1 (en) * 2009-03-04 2010-09-10 Airbus Operations Gmbh Wing of an aircraft and assembly of a wing comprising a device for influencing a flow
US9079657B2 (en) 2009-03-04 2015-07-14 Airbus Operations Gmbh Wing of an aircraft and assembly of a wing comprising a device for influencing a flow
DE102019130596A1 (en) * 2019-11-13 2021-05-20 Deutsches Zentrum für Luft- und Raumfahrt e.V. Method for suppressing aeroelastic instabilities on transonic operating aircraft and aircraft with devices for carrying out the method by means of direct intervention in the aerodynamic boundary layer
DE102019130596B4 (en) 2019-11-13 2022-10-27 Deutsches Zentrum für Luft- und Raumfahrt e.V. Method for suppressing aeroelastic instabilities on transonic aircraft and aircraft with devices for carrying out the method by direct intervention in the aerodynamic boundary layer

Also Published As

Publication number Publication date
DE4207103C1 (en) 1993-09-16
EP0558904B1 (en) 1995-11-22
US5335885A (en) 1994-08-09

Similar Documents

Publication Publication Date Title
EP0558904B1 (en) Aircraft wing with supercritical profile
EP1149761B1 (en) Noise damping device for aircraft wings
DE2254888A1 (en) WINGS FOR SUBSOUND SPEED
EP0071616B1 (en) Lift-generating element, particularly for aircraft wings and blades of turbomachines, such as air screws and ship propellers
DE3342421A1 (en) METHOD FOR THE STABILIZING INFLUENCE OF DETACHED LAMINARY BORDER LAYERS
DE1628237A1 (en) Supersound grille
DE3534268C2 (en)
CH422532A (en) Airplane fan inlet duct for vertical takeoff and landing
DE69910521T2 (en) Method for reducing the wave resistance of an aircraft
EP2429895B1 (en) Fairing for a high-lift device
EP0126199B1 (en) Aircraft wing
DE2949133C2 (en) Airfoil with supercritical profiling
DE102004013645A1 (en) Axial compressor cascade blade design curves outside and inside of blades to give under- and overpressure respectively and specified compensatory supersonic constant speed region for performance boost.
DE102019008248B4 (en) FLOW PROFILE AND MECHANICAL MACHINE EQUIPPED THEREFROM
DE4442319C1 (en) Aircraft wing designed for speeds in high subsonic range
DE1809593B2 (en) Device for influencing the eddies on the upper side of wings
EP1506922B1 (en) Method and aerodynamic body with device for reducing the drag of a limited supersonic flow region
DE10305973B3 (en) Loss reduction device for compression shock includes at least one shock-inducing body spaced out from surface in supersonic flow region
DE102019130596B4 (en) Method for suppressing aeroelastic instabilities on transonic aircraft and aircraft with devices for carrying out the method by direct intervention in the aerodynamic boundary layer
DE3836673A1 (en) Method and arrangement for reducing the induced drag of a wing
DE1286405B (en) Airplane with a helicopter rotor
DE102015220671B4 (en) Aerodynamic body, its use and method of aspirating a portion of a fluid
DE938645C (en) Aircraft structure or the like.
DE102021115604A1 (en) Anti-wing device for wings
DE102010045080B4 (en) airfoil

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): FR GB IT NL

17P Request for examination filed

Effective date: 19930813

17Q First examination report despatched

Effective date: 19941111

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: DAIMLER-BENZ AEROSPACE AIRBUS GESELLSCHAFT MIT BES

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): FR GB IT NL

GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)

Effective date: 19960116

ITF It: translation for a ep patent filed

Owner name: INVENTION S.N.C.

NLT2 Nl: modifications (of names), taken from the european patent patent bulletin

Owner name: DAIMLER-BENZ AEROSPACE AIRBUS GESELLSCHAFT MIT BES

ET Fr: translation filed
PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed
REG Reference to a national code

Ref country code: FR

Ref legal event code: CD

REG Reference to a national code

Ref country code: GB

Ref legal event code: IF02

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: IT

Payment date: 20100127

Year of fee payment: 18

Ref country code: FR

Payment date: 20100223

Year of fee payment: 18

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20100121

Year of fee payment: 18

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: NL

Payment date: 20100118

Year of fee payment: 18

REG Reference to a national code

Ref country code: NL

Ref legal event code: V1

Effective date: 20110801

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 20110122

REG Reference to a national code

Ref country code: FR

Ref legal event code: ST

Effective date: 20110930

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20110131

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20110122

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20110122

Ref country code: NL

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20110801